This invention relates to attitude control systems for spacecraft, and more particularly to systems which reduce attitude jitter.
FIG. 1 is a simplified block diagram of a spacecraft, designated generally as 10, orbiting about a heavenly body 12, which may be Earth. Spacecraft 10 includes a body 14 oriented relative to roll, pitch and yaw axes 16, 18, and 20, respectively. Yaw axis 20 is directed toward earth, roll axis 16 points in the orbital direction, and pitch axis 18 is normal to the orbital plane. Satellite body 14 supports solar panels 22a and 22b as known for producing electrical energy for powering spacecraft attitude control systems and the electrical portions (not illustrated) of housekeeping systems, torquers such as magnetic torquers and arc jet thrusters, if used, and for the payload. An earth sensing arrangement (ESA) illustrated as 24 is mounted on spacecraft body 14 for producing signals representative of the roll and/or pitch attitude of body 14. A sun sensor illustrated as 44 produces yaw-representative signals at certain times of day, as is well known in the art.
FIG. 2 is a simplified block diagram of a prior art attitude control arrangement usable in the satellite of FIG. 1. Elements of FIG. 2 corresponding to those of FIG. 1 are designated by like reference numerals. In FIG. 2, earth sensor assembly 24 produces roll and attitude signals which are applied to the inverting (-) input ports of roll and pitch, respectively, summing circuits 30 and 32. Noninverting (+) input ports of summing circuits 30 and 32 are coupled to receive roll and pitch, respectively, attitude commands from a source of attitude commands illustrated as a block 34. As known, block 34 may include a memory in which commands are stored for application at predetermined times, and may include a communication channel by which the stored attitude command information may be updated. Summing circuits 30 and 32 subtract the sensed attitude from the commanded attitude to produce roll and pitch error signals, respectively, which are applied to proportional-derivative (PD) wheel controllers 36 and 38, respectively. The roll and pitch attitude signals produced by ESA 24 are also applied to differentiating (d/dt) blocks 26 and 28, respectively. The differentiated signals produced at the outputs of blocks 26 and 28 represent roll and pitch rates, respectively. The roll and pitch rate signals generated at the outputs of differentiating blocks 26 and 28, respectively, are applied to proportional-derivative wheel controllers 36 and 38, respectively. Proportional-derivative controllers 36 and 38 each calculate a linear combination of proportional and derivative signals, and more particularly, PD controller 36 multiplies roll attitude rate from block 26 by a roll attitude rate gain, multiplies roll attitude error from block 30 by a roll position gain, and sums together the two products to produce roll torque command signals for driving a roll wheel drive circuit 40a for control of a roll wheel represented by block 40b. PD controller 38 multiplies pitch attitude rate from block 28 by a pitch attitude rate gain, multiplies pitch attitude error from summing circuit 32 by a pitch position gain, and sums together the two products to produce pitch torque command signals for driving a pitch wheel drive circuit 42a for control of a pitch wheel represented by a block 42b.
A sun sensor illustrated as a block 44 in FIG. 2 at least periodically produces signals representative of the spacecraft yaw attitude. The yaw attitude signals are applied to the inverting input of a summing circuit 48, in which the sensed yaw attitude is subtracted from the commanded yaw attitude produced in block 34, to produce a yaw attitude error signal for application to a yaw proportional-derivative controller 50. The yaw attitude signals from sun sensor 44 are also applied to a differentiator 46 to produce yaw rate signals, which are applied as a second input to PD controller 50. Controller 50 sums together the yaw attitude error and the yaw rate to produce torque command signals for application to a yaw wheel drive 52a and to a yaw wheel 52. The axis of rotation of wheel 52b is at least parallel to yaw axis 20 of FIG. 1.
Those persons skilled in the art to which the invention pertains will recognize that, if the roll, pitch or yaw wheels, or any of them are skewed relative to the spacecraft axes, the commanded torques must be distributed among the wheels to achieve the desired attitude control about a particular axis.
The signals produced by the ESA and the sun sensor may be noisy, either because of actual noise inherent in the sensor, or because of poor viewing angle, or possibly because of unwanted objects in the field of view, and also can be noisy because of poor resolution attributable to an insufficient number of quantizing levels. The differentiation or time derivative of the noisy signal in the differentiators 26, 28 and/or 46 amplifies high frequency noise. The amplified high frequency noise becomes part of the control signal driving the wheel.
FIG. 3a illustrates a torque-amplitude vs. time plot 310 for the prior art arrangement of FIG. 2. As illustrated, substantial torque variations occur, even in regions in which the attitude error is essentially constant. The peak-to-peak excursions of the torque are about 0.4 inch-pounds (in-lbs). These large, high frequency torque components represent energy applied to the wheels. As a result, electrical energy is consumed by the wheels without producing a net error correction, the wheels are subjected to continuous forces, thereby tending to wear their bearings. If chemical thrusters are used instead of wheels, propellant is wasted by the frequent operation of the thrusters. The frequent torquing by use of wheels or thrusters necessarily results in pointing jitter.
One possible solution to the problem of jitter in the arrangement of FIG. 2 is to use improved, low noise or high resolution sensors, or to provide a long-life gyroscope for use during slewing maneuvers. However, space-qualified high-resolution sensors and gyroscopes are extremely expensive, and may be prohibitively so for commercial spacecraft.
Another possible solution would be to low pass filter the sensor signal in the arrangement of FIG. 2, to thereby reject high-frequency sensor components. However, this has a major disadvantage. If the pointing of the spacecraft is to be changed, as might be the case if it were desired to direct a sensing instrument toward a different part of the earth, or a telescope toward a different part of the sky, a step command would be generated by attitude command block 34 and applied to one or more of summing circuits 30, 32 and 48. This would initiate a fast slew toward the new attitude as the wheels were torqued. However, the filtered sensor signals could not produce high frequency sensed attitude signals, whereby the sensed attitude signals would lag the actual attitude. Consequently, the actual attitude of the spacecraft would overshoot the commanded attitude, and "hunt" toward the commanded attitude. Depending upon the damping of the system, this might or might not converge, but in any case is time-consuming and expends more fuel or electrical energy than is necessary to accomplish the desired attitude change.